Combustor with annular bluff body

ABSTRACT

The present invention relates to a gas turbine combustor comprising: a flow sleeve; a combustion liner located at least partially within the flow sleeve thereby creating a main passage between the flow sleeve and the combustor liner; a dome located forward of the flow sleeve and encompassing at least a part of the combustion liner, the dome having a substantially rounded head end thereby forming a turning passage between the liner and the head end; and a swirler wall aligned along a centerline of the combustor, the swirler wall projecting into the liner, wherein the swirler wall and the rounded head end are connected, wherein the connection forms an annular end face.

TECHNICAL FIELD

The present invention relates generally to a system and method forimproving combustion stability in a gas turbine combustor.

BACKGROUND

In an effort to reduce the amount of pollution emissions fromgas-powered turbines, governmental agencies have enacted numerousregulations requiring reductions in the amount of oxides of nitrogen(NOx) and carbon monoxide (CO). Lower combustion emissions can often beattributed to a more efficient combustion process, with specific regardto fuel injector location and mixing effectiveness.

Early combustion systems utilized diffusion type nozzles, where fuel ismixed with air external to the fuel nozzle by diffusion, proximate theflame zone. Diffusion type nozzles have been known to produce highemissions due to the fact that the fuel and air burn stoichiometricallyat high temperature to maintain adequate combustor stability and lowcombustion dynamics.

An enhancement in combustion technology is the utilization of premixing,such that the fuel and air mix prior to combustion to form a homogeneousmixture that burns at a lower temperature than a diffusion type flameand produces lower NOx emissions. Premixing fuel and air together beforecombustion allows for the fuel and air to form a more homogeneousmixture, which for a given combustor exit temperature will burn at lowerpeak temperatures, resulting in lower emissions. Example of such a gasturbine flamesheet combustion system with reduced emissions and improvedflame stability at multiple load conditions is disclosed in US patentapplication US2004/0211186A1

While the combustors of the prior art have improved emissions levels andability to operate at reduced load settings, thermoacoustics of theflamesheet combustors could still lead to instability modes (such aspulsation), which could restrict the operation window. Additionally,aerodynamics of the burner allows occasional flame attachment in themixing zone under certain circumstances, causing flashback andoverheating risk. Furthermore, current fuel staging strategies couldcause asymmetrical heat load on the combustor liner, which could lead tocreep problems.

In addition, measure which help against pulsation, as for example thestaging of 1/3-2/3 groups in the main fuel supply can lead toasymmetrical liner heat loading, as well as to non-uniformities in thecombustor exit temperature profile.

What is intended is a system that can provide further flame stabilitywhile also reducing thermoacoustic instabilities which can enlarge theoperation window available of the current combustor designs. Theembodiments described below are intended to widen the operation windowbeyond the currently available range, without sacrificing the lowemission values.

SUMMARY OF THE INVENTION

It is one object of the present invention to provide a combustor withfurther improved stability and improved thermoacoustics characteristics.

The above and other objects of the invention are achieved by a gasturbine combustor comprising a flow sleeve, a combustion liner locatedat least partially within the flow sleeve thereby creating a mainpassage between the flow sleeve and the combustor liner, and a domelocated forward of the flow sleeve and encompassing at least a part ofthe combustion liner, the dome having a substantially rounded head endthereby forming a turning passage between the liner and the head end,and a swirler wall aligned along a centerline of the combustor, theswirler wall projecting into a space delimited by the liner, wherein theswirler wall and the rounded head end are connected, and wherein theconnection forms an annular end face.

According to one embodiment of the present invention, the combustorfurther comprises a center body positioned along the centerline andextending into the space delimited by the swirler wall, thereby forminga pilot passage between the swirler wall and the center body.

According to yet another embodiment of the present invention, width ofthe pilot channel is substantially constant along the length of thepilot channel.

According to another embodiment of the present invention, an area of theannular end face is 1.5 times to 5 times larger than an area of a crosssection of the pilot passage.

According to yet another embodiment of the present invention, a fuellance is arranged in the center body.

According to another embodiment of the present invention, the combustorfurther comprises a substantially cylindrical extension extending from aradially inner end of the rounded head end or the end face into theliner, wherein the extension is aligned with the centerline of thecombustor. According to yet another embodiment of the present invention,the extension has substantially constant radius along the centerline ofthe liner, and/or the thickness of the extension is substantially equalto the thickness of the rounded head end.

According to another embodiment of the present invention, a recessdelimited by the central body, the annular end face and the rounded headend comprises a Helmholtz damper or/and means for pilot oil injection.

According to yet another embodiment of the present invention, thecombustion liner comprises a ring shaped rounded lip section and acurved middle section adapted to create a flame stabilization zoneduring operation. According to another embodiment of the presentinvention the lip section comprises a Helmholtz damper and/or liquidfuel injection means.

According to another embodiment of the present invention, the pilotpassage comprises a pilot swirler in fluid communication with at leastone pilot fuel injector, and the pilot swirler is an axial swirler or aradial swirler. According to another embodiment of the presentinvention, the main passage or the turning passage comprises a mainswirler in a fluid communication with at least one main fuel injector,and wherein the main swirler is an axial swirler or a radial swirler.

According to another embodiment of the present invention, the swirlerwall is a part of a conical burner (e.g. EV burner or AEV burner).

The present application also provides for a gas turbine comprising thecombustor described above.

In addition, the present application also provides for a method foroperating the gas turbine combustor. The method comprising: supplying afirst stream of fuel into the pilot channel or conical burner (e.g. EVburner or AEV burner) to mix with the first flow of air, and feeding theresulting first mixture into the combustion zone for providing pilotflame; supplying a second flow of air into the main passage; supplying asecond stream of fuel into the main passage or turning passage to mixwith the second flow of air, and feeding the resulting second mixtureinto the combustion zone for providing a main flame.

Additional advantages and features of the present invention will be setforth in part in a description which follows, and in part will becomeapparent to those skilled in the art upon examination of the following,or may be learned from practice of the invention. The instant inventionwill now be described with particular reference to the accompanyingdrawings.

BRIEF DESCRIPTION OF DRAWINGS

Preferred embodiments of the invention are described in the followingwith reference to the drawings, which are for the purpose ofillustrating the present preferred embodiments of the invention and notfor the purpose of limiting the same. In the drawings,

FIG. 1 shows a cross section view of a gas turbine combustion system ofthe prior art.

FIG. 2a shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention.

FIG. 2b shows an end view of a gas turbine combustor in accordance withan embodiment of the present invention.

FIG. 2c shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention schematicallyindicating flame fronts during operation.

FIG. 3a shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention.

FIG. 3b shows an end view of a gas turbine combustor in accordance withan embodiment of the present invention.

FIG. 4a shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention.

FIG. 4b shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention.

FIG. 5 shows a cross section view of a gas turbine combustor inaccordance with an embodiment of the present invention schematicallyindicating recirculation zones used for further flame stabilization.

FIGS. 6a, 6b, 6c show a cross section view of a part of a gas turbinecombustor in accordance with embodiments of the present invention.

FIG. 7a shows cross section view of a part of a gas turbine combustorcomprising EV burner in accordance with embodiments of the presentinvention.

FIG. 7b shows cross section view of a part of a gas turbine combustorcomprising AEV burner in accordance with embodiments of the presentinvention.

FIG. 8a shows a perspective view of a part of EV burner

FIG. 8b shows a cross section view of a part of AEV burner.

DETAILED DESCRIPTION OF THE DRAWINGS

An example of a premixing flamesheet combustor 100 for a gas turbine ofthe prior art is shown in FIG. 1. The combustor 100 is a type of reverseflow premixing combustor utilizing a pilot nozzle 102, a radial inflowmixer 104, and a plurality of main stage mixers 108. The pilot portionof the combustor 100 is separated from the main stage combustion area bya center divider portion 110. The center divider portion 110 separatesthe fuel injected by the pilot nozzle 102 from the fuel injected by themain stage mixers 108. Correspondingly the air entering through the mainand the pilot burner is separated by the divider 110. A flame front 120,which might occur for an off-design case, is shown schematicallyindicating interaction of pilot and main flame, which might causethermoacoustic instabilities.

FIG. 2a shows a cross section view of a gas turbine combustor 200 inaccordance with an embodiment of the present invention. The combustor200 comprising a flow sleeve 202, a combustion liner 204 located atleast partially within the flow sleeve 202 thereby creating a mainpassage 206 between the flow sleeve 204 and the combustor liner 204. Thecombustor 200 also comprises a dome 208 located forward of the flowsleeve and encompassing at least a part of the combustion liner 204. Thedome 208 has a substantially rounded head end 210 thereby forming aturning passage 212 between the liner 204 and the head end 210. Thecompressor 200 comprises also a swirler wall 214 aligned along acenterline 216 of the combustor 200, wherein the swirler wall 214 isprojecting into the liner 204. The swirler wall 214 and the rounded headend 210 are connected, wherein the connection forms an annular end face218. The structure and thickness of the end face 218 can vary, and inone embodiment the end face 218 is a thin plate, for example a sheetmetal plate. In one embodiment the end face 218 has a flat surfacesubstantially perpendicular to the centerline 216. In one embodiment ofthe present invention, the end face 218 is cooled via effusion and/orimpingement cooling.

In one embodiment according to the invention, the combustor 200 furthercomprises a center body 220 positioned along the centerline 216 andextending into the space delimited by the swirler wall 214. The swirlerwall 214 and the center body 220 form a pilot passage 222. The centerbody comprises a front surface 226 which can have different shapes,depending on the combustor design, such as bluff body shape. The widthof the pilot channel 222 can vary, and preferably is substantiallyconstant along the length of the pilot channel 222. The center body 220could also comprise a fuel lance 608 (shown in FIG. 6b ) to create acentral pilot flame.

FIG. 2b shows an end view of a gas turbine combustor in accordance withan embodiment of the present invention. The cross sections of differentcomponents are shown as a generally cylindrical, but they can have othershapes such as oval or elongated. An area of the annular end face 218can vary in respect of the size of the other components of the combustor200. In one preferred and non-limiting example, the area of the annularend face 218 is 1.5 times to 5 times larger than an area of a crosssection of the pilot passage 222.

The combustor 200 according to the invention in one embodiment cancomprise main fuel supply 234, pilot fuel supply 230, main swirler withinjectors 232 and pilot swirler with injectors 228 to create a pilotflame and a main flame during an operation of the combustor. FIG. 2cshows schematically flame fronts, inside a combustion zone 250, createdduring operation of the combustor 200 according to the presentinvention. Contrary to the prior art (FIG. 1) where the pilot flame andthe main flame interacts, in the embodiment according to the invention amain flame 260 and a pilot flame 262 are clearly separated due to theadvantageous design of the combustor 200 according to the invention.

FIG. 3a shows a cross section view of a gas turbine combustor 200 inaccordance with another embodiment of the present invention whichfurther comprises a substantially cylindrical extension 240 extendingfrom a radially inner end of the rounded head end 210 into the liner220. In an alternative embodiment, the extension 240 is extending fromthe end face 218. The extension is substantially aligned with thecenterline 216 of the combustor 200. The extension 240 can vary in size,length, radius and width depending on operating parameters of thecombustor 200. In one embodiment, the extension 240 is cylindrical andit has substantially constant radius along the centerline 216 of theliner. In one embodiment according to the invention the extension 240and head end 210 have substantially same thickness. The extension 240and head end 210 could be made as two separate pieces or they can bemade of a single piece of material. In one embodiment, the extension 240and head end 210 are made of a sheet metal. The cooling of the extension240 may be done by near wall cooling using channels in axial direction.

FIG. 3b shows an end view of a gas turbine combustor in accordance withan embodiment of the present invention shown in FIG. 3a . In oneembodiment, an average thickness of extension 240 is smaller thanaverage thickness of a cross section of the end face 218.

FIG. 4a shows a cross section view of a gas turbine combustor 200 inaccordance with yet another embodiment according to the presentinvention wherein the combustion liner 204 comprises a ring shapedrounded lip section 420 and a curved middle section 430. The liner 204according to this embodiment could also comprise cooling holes 440. Inthis embodiment, the rounded lip section 420 is substantially hollow.FIG. 4b shows an alternative embodiment, wherein the rounded lip sectionis made of thin material, substantially of the same thickness as themain portion of the liner 204, for example of a sheet metal. In thisway, reducing the thickness of the rounded lip 420, there isadvantageously more room for a stabilization zone.

The embodiment comprising the ring shaped rounded lip section 420 andthe curved middle section 430 is adapted to create an additional outermain flame stabilization zone 510 during operation as shown in FIG. 5.FIG. 5 also shows a central pilot stabilization zone 530 and an outerpilot stabilization zone 520 created during operation of combustor 200according to the invention. The extension 420 advantageously makespossible effective separation of two pilot stabilization zones 520 and530.

FIGS. 6a, 6b and 6c show additional embodiments of the presentinvention. The lip section of the liner could comprise a Helmholtzdamper 612 and/or liquid fuel injection means 606. A recess 242delimited by the central body 214, the annular end face 218 and therounded head end 210 could comprises a Helmholtz damper 610 or/and ameans for pilot oil injection 604. In general, Helmholtz damper isdesigned according to an individually determined or predetermineddamping requirement against the thermoacoustic oscillation frequenciesoccurring in the combustion chamber. The Helmholtz damper comprises adamper volume, a neck and a cooling channel. The pilot swirlers (228,618) and the main swirlers (232, 620) in general could be axial orradial swirlers. In addition, the combustor 200 may comprise additionalHelmholtz damper 602 and the fuel lance 608, both inside the center body220, as shown in FIG. 6a and FIG. 6 b.

The combustor 200 according to the invention could comprise a conicalburner 702,704 device instead of the center body 220. Examples of theseembodiments are shown in FIGS. 7a and 7b , including EV burner(environmental burner from Alstom, disclosed in EP0321809) and AEVburner (advanced environmental burner from Alstom, disclosed inEP0704657) respectively. In these embodiments, the swirler wall 214 is apart of the conical burner 702,704.

FIG. 8a shows part of EV burner 702 wherein a conical column 5 of liquidfuel is formed in the interior 14 of the burner 702, which column widensin the direction of flow and is surrounded by a rotating stream 15 ofcombustion air which flows tangentially into the burner. Ignition of themixture takes place at the burner outlet, a backflow zone 6 forming inthe region of the burner outlet. The burner itself consists of at leasttwo hollow part-cone bodies 1, 2 which are superposed on one another andhave a cone angle increasing in the direction of flow. The part-conebodies 1, 2 are mutually offset. A nozzle 3 placed at the burner headensures injection of the liquid fuel 2 into the interior 14 of theburner. In one embodiment of the present invention, in the combustor 200according to the invention, part cone body 1 of EV burner 702corresponds to the swirler wall 212.

FIG. 8b shows part of AEV burner 704 comprising of at least part of theEV burner 702 and a mixing tube 802. The mixing tube comprises a tube804. In one embodiment of the present invention, in the combustor 200according to the invention, the tube 804 of AEV burner 704 correspondsto the swirler wall 212.

It should be apparent that the foregoing relates only to the preferredembodiments of the present application and that numerous changes andmodifications may be made herein by one of ordinary skill in the artwithout departing from the general spirit and scope of the invention asdefined by the following claims.

LIST OF DESIGNATIONS

-   1,2 Part cone bodies-   3 Nozzle-   5 Conical column-   6 Backflow zone-   14 Interior of a burner-   15 Rotating stream-   100 Combustor-   102 Pilot nozzle-   104 Radial inflow mixer-   108 Main stage mixer-   110 Divider-   120 Flame front-   200 Combustor-   202 Flow sleeve-   204 Combustion liner-   206 Main passage-   208 Dome-   210 Head end-   212 Turning passage-   214 Swirler wall-   216 Combustor centerline-   218 End face-   220 Center body-   222 Pilot passage-   226 Center body front surface-   228 Pilot swirler with injectors-   230 Pilot fuel supply-   232 Main swirler with injectors-   234 Main fuel supply-   240 Extension-   242 Recess-   250 Combustion zone-   260 Main flame-   262 Pilot flame-   420 Lip section-   430 Curved middle section-   440 Cooling holes-   510 Main flame stabilization zone-   520 Outer pilot stabilization zone-   530 Central pilot stabilization zone-   602 Helmholtz damper-   604 Pilot oil injection-   606 Oil injection-   608 Fuel lance-   610 Helmholtz damper-   612 Helmholtz damper-   614 Fuel injector-   618 Pilot swirler-   620 Main swirler-   622 Fuel injector-   702 EV burner-   704 AEV burner-   802 Mixing section-   804 Tube

1. A gas turbine combustor comprising: a flow sleeve; a combustion linerlocated at least partially within the flow sleeve thereby creating amain passage between the flow sleeve and the combustor liner; a domelocated forward of the flow sleeve and encompassing at least a part ofthe combustion liner, the dome having a substantially rounded head endthereby forming a turning passage between the liner and the head end;and a swirler wall aligned along a centerline of the combustor, theswirler wall projecting into a space delimited by the liner, wherein theswirler wall and the rounded head end are connected, wherein theconnection forms an annular end face.
 2. The combustor of claim 1further comprising a center body positioned along the centerline andextending into the space delimited by the swirler wall, thereby forminga pilot passage between the swirler wall and the center body.
 3. Thecombustor of claim 1 wherein a width of the pilot channel issubstantially constant along the length of the pilot channel.
 4. Thecombustor of claim 1, wherein an area of the annular end face is 1.5times to 5 times larger than an area of a cross section of the pilotpassage.
 5. The combustor of claim 1, wherein a fuel lance is arrangedin the center body.
 6. The combustor of claim 1 further comprising asubstantially cylindrical extension extending from a radially inner endof the rounded head end or the end face into the liner, wherein theextension is aligned with the centerline of the combustor.
 7. Thecombustor of claim 8, wherein the extension has substantially constantradius along the centerline of the liner, and/or wherein the thicknessof the extension is substantially equal to the thickness of the roundedhead end.
 8. The combustor of claim 1, wherein a recess delimited by thecentral body, the annular end face and the rounded head end comprises aHelmholtz damper or/and a means for pilot oil injection.
 9. Thecombustor of claim 1, wherein the combustion liner comprises a ringshaped rounded lip section and a curved middle section adapted to createa flame stabilization zone during operation.
 10. The combustor of claim8, wherein the lip section comprises a Helmholtz damper and/or liquidfuel injection means.
 11. The combustor of claim 1, wherein the pilotpassage comprises a pilot swirler in fluid communication with at leastone pilot fuel injector, and wherein the pilot swirler is an axialswirler or a radial swirler.
 12. The combustor of claim 1, wherein themain passage or the turning passage comprises a main swirler in a fluidcommunication with at least one main fuel injector, and wherein the mainswirler is an axial swirler or a radial swirler.
 13. The combustor ofclaim 1, wherein the swirler wall is a part of a conical burner.
 14. Agas turbine comprising the combustor according to claim
 1. 15. A methodfor operating the gas turbine combustor according to claim 1, the methodcomprising: supplying a first stream of fuel into the pilot channel orthe conical burner, and feeding the resulting first mixture into thecombustion zone for providing pilot flame; supplying a second flow ofair into the main passage; supplying a second stream of fuel into themain passage or turning passage to mix with the second flow of air, andfeeding the resulting second mixture into the combustion zone forproviding a main flame.